A conventional midframe design for a can-annular gas turbine engine is discussed in U.S. Pat. No. 7,721,547 (“'547 Patent”), assigned to the assignee of the present invention, which is incorporated by reference herein. FIG. 1 of the '547 Patent is reproduced as FIG. 1 herein, and illustrates a cross-section through a midframe portion 13 of a conventional can-annular gas turbine engine 10. The major components of the gas turbine engine 10 are a compressor section 12, a combustion section 16 and a turbine section 48. A rotor assembly 17 is centrally located and extends through the three sections. In operation, the compressor section 12 receives air through an intake (not shown) and compresses it. The compressed air flow 11 passes from the compressor section 12 to an axial diffuser 14, after which the air flow 11 enters a chamber 15 within a casing 19, where the total air flow 11 is separated and enters one of multiple combustor heads 18 of the can-annular combustion section 16 that encircle the rotor assembly 17 in an annular configuration.
As illustrated in FIG. 1, the compressor section 12 includes cylinders 27,29 that enclose alternating rows of stationary vanes 23 and rotating blades 25. The stationary vanes 23 can be affixed to the cylinder 27 while the rotating blades 25 can be mounted to the rotor assembly 17 for rotation with the rotor assembly 17. The stationary vanes 23 include a last stationary vane 26 and an outlet guide vane 28 positioned adjacent to an outlet of the compressor section 12. Additionally, the rotating blades 25 include a last stage blade 24 positioned upstream from the last stationary vane 26 and the outlet guide vane 28. The last stationary vane 26 and outlet guide vane 28 are used to remove an absolute tangential swirl angle (measured in an absolute reference frame with respect to the longitudinal direction) of the air flow 11 coming off the last stage blade 24.
As further illustrated in FIG. 1, load-bearing struts 30 are provided to support a shaft cover 32 of the rotor assembly 17 at the casing 19 of the combustion section 16. As appreciated by one of skill in the art, one strut 30 may be provided per each one to four combustor heads 18. As illustrated in FIG. 1, the axial diffuser 14 includes an inner cone 36 and an outer cone 34 and the cross-sectional area between the inner and outer cones 36,34 increases in the longitudinal direction 68, such that the air flow 11 expands and decelerates through the diffuser 14, thereby converting velocity head into pressure head. As illustrated in FIG. 1, the strut 30 is attached between a shaft cover 32 to the outer cone 34 of the axial diffuser 14, and thus the casing 19 of the combustion section 16 supports the strut 30 at the shaft cover 32.
As further illustrated in FIG. 1, a rotor-cooling extraction pipe 38 is provided, which extracts compressed air from the chamber 15 and passes the compressed air into a cooler 42. The cooled air passes from the cooler 42 and through rotor-cooling injection pipes 40 that are positioned within the chamber 15 and direct the cooled air below the shaft cover 32, to cool the rotating components of the engine.
Another portion of the engine needing cooling is a turn in the transition 20 at an inlet to the turbine section 48, which typically experiences an especially high heat flux during an operation of the gas turbine engine 10. In order to cool a rear end 54 of the transition 20 during operation of the gas turbine 10, a portion 58 of the air flow 11 entering the chamber 15 makes contact with the rear end 54 of the transition 20 proximate the highest heat flux region in order to cool the rear end 54 of the transition 20 using thermal convection.
FIG. 8 of the '547 Patent is reproduced herein as FIG. 2, and illustrates a “trans-vane” transition 20′ which improves upon the transition 20 of FIG. 1. FIG. 2 illustrates a top down radial view of the midframe portion 13′ of the gas turbine engine 10′ including the combustion section 16′ and a first stage turbine blade array 49′ of the turbine section 48′ located downstream from the combustion section 16′, with the trans-vane transition 20′ located therebetween. The midframe portion 13′ of FIG. 2 includes a compressor section (not shown) similar to the compressor section 12 of FIG. 1. A first stage housing encloses the first stage turbine blade array 49′ and includes a blade ring 51′. An upstream side 53′ of the blade ring 51′ is preferably adapted to couple to a transition outlet 55′. The trans-vane transition 20′ includes a transition duct body 60′ with an inlet 62′ to receive a gas flow exhausted from the combustor section 16′ and the outlet 55′ to discharge a gas flow toward the first stage blade array 49′ with an internal passage 66′ therebetween. The outlet 55′ is offset from the inlet 62′ in the three coordinate directions—in the radial direction (in/out of the figure), the longitudinal direction 68 and the tangential direction 70′. The gas flow discharged from the outlet 55′ is angled in the tangential direction 70′ within an absolute reference frame, relative to the longitudinal direction 68 as depicted by the arrow 72′, as required by the first stage turbine blade array 49′. A brief discussion will be provided of the absolute and relative reference frames of the midframe portion 13′, as well as how the velocity vector of an air flow exiting the compressor and entering the turbine 48′ of the gas turbine engine 10′ is represented in each of those reference frames. As discussed below, the velocity vector of an air flow exiting the compressor or entering the turbine is measured in an absolute reference frame, with respect to a longitudinal direction 68 along the longitudinal axis 75 of the gas turbine engine. FIG. 3 illustrates a top down radial view of the last stage blade 24 of the compressor section 12 of the gas turbine engine 10′ and the first stage blade 49′ of the turbine 48′ of the midframe portion 13′, separated along a longitudinal axis 75 of the conventional gas turbine engine 10′ of FIG. 2 oriented along the longitudinal direction 68. An outgoing air flow off the last stage blade 24 is oriented in the (relative) reference frame of the last stage blade 24 along a relative outgoing velocity vector 76. During an operation of the compressor section 12, the last stage blade 24 rotates around the longitudinal axis 75 with a blade velocity vector 78 that is oriented perpendicular to the longitudinal axis 75. In order to determine the velocity vector of the outgoing air flow off the last stage blade 24 in an absolute reference frame, the blade velocity vector 78 is added to the relative outgoing velocity vector 76, resulting in an absolute outgoing velocity vector 80 that is angled in the tangential direction 70 by an angle 82, relative to the longitudinal axis 75 oriented in the longitudinal direction 68. In an exemplary embodiment, the angle 82 is approximately 45 degrees. Accordingly, the absolute outgoing velocity vector 80 of the outgoing air flow off the last stage blade 24 is oriented approximately 45 degrees in the tangential direction 70, relative to the longitudinal axis 75 oriented in the longitudinal direction 68. The last stage vanes 26,28 of the conventional midframe portion 13′ are configured to reduce the angle 82 of the absolute outgoing velocity vector 80 from 45 degrees to approximately 0 degrees, to align the air flow along the longitudinal axis 75. However, as discussed below, the embodiments of the present invention do not utilize the last stage vanes, and thus utilize the initial angle 82 of the absolute outgoing velocity vector 80 off the last stage blade 24. FIG. 3 also illustrates an incoming air flow to the first stage blade 49′ of the turbine 48′ illustrated in FIG. 2. In order to maximize the effectiveness of the turbine 48′, the incoming air flow is oriented in the (relative) reference frame of the first stage blade 49′ along a relative incoming velocity vector 84. During an operation of the turbine 48′, the first stage blade 49′ rotates around the longitudinal axis 75 with a blade velocity 86 that is oriented perpendicular to the longitudinal axis 75. In order to determine the velocity vector of the incoming air flow in the absolute reference frame, the blade velocity vector 86 is added to the relative incoming velocity vector 84, resulting in an absolute incoming velocity vector 88 that is angled in the tangential direction 70 by an angle 90, relative to the longitudinal direction 68. In an exemplary embodiment, the angle 90 is approximately 70 degrees. Accordingly, the absolute incoming velocity vector 88 of the incoming air flow onto the first stage blade 49′ of the turbine 48′ is oriented approximately 70 degrees in the tangential direction 70, relative to the longitudinal direction 68. In contrast with the transition 20′ of FIG. 2, the transition 20 illustrated in FIG. 1 discharges a gas flow to the turbine section 48 with an offset in only the radial direction and the longitudinal direction 68, and thus the gas flow is not angled in the tangential direction relative to the longitudinal direction 68. Since the first stage turbine blade array 49 of the turbine section 48 requires an incoming gas flow that is angled in the tangential direction relative to the longitudinal direction 68, the turbine section 48 of FIG. 1 includes a first stage vane 74, to introduce an offset in the tangential direction for the gas flow discharged from the transition 20. However, by implementing the trans-vane design in the transition 20′, the gas flow is discharged from the outlet 55′ at the necessary angle 90 in the tangential direction 70 relative to the longitudinal direction 68 to accommodate the first stage turbine blade array 49′, and thus the first stage vanes 74 are not needed. In the '547 Patent, the inventors made various improvements to the midframe portion of the gas turbine engine, downstream of the combustion section, to enhance the operating efficiency of the gas turbine engine. In the present invention, the present inventors make various improvements to the midframe portion of the gas turbine engine, upstream of the combustion section, to also enhance the operating efficiency and/or cost efficiency of the gas turbine engine.